U.S. patent application number 15/830450 was filed with the patent office on 2018-03-29 for blade outer air seal surface.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Patrick D. Couture, Paul M. Lutjen.
Application Number | 20180085880 15/830450 |
Document ID | / |
Family ID | 51297538 |
Filed Date | 2018-03-29 |
United States Patent
Application |
20180085880 |
Kind Code |
A1 |
Lutjen; Paul M. ; et
al. |
March 29, 2018 |
BLADE OUTER AIR SEAL SURFACE
Abstract
A blade outer air seal for a gas turbine engine having a surface
that is eccentric with respect to the engine rotation centerline,
and a method for creating same, are disclosed. Also, a method for
grinding a work piece having nominal curvature defined by a work
piece curvature centerline is disclosed, comprising the steps of:
a) determining a desired surface profile for the work piece; b)
providing a rotating grinding surface having a grinding rotation
centerline; c) offsetting the grinding rotation centerline from the
work piece curvature centerline; and d) applying the rotating
grinding surface to the work piece while rotating the rotating
grinding surface about the grinding rotation centerline to create
the desired surface profile.
Inventors: |
Lutjen; Paul M.;
(Kennebunkport, ME) ; Couture; Patrick D.;
(Tolland, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
51297538 |
Appl. No.: |
15/830450 |
Filed: |
December 4, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
14176669 |
Feb 10, 2014 |
9833869 |
|
|
15830450 |
|
|
|
|
61763231 |
Feb 11, 2013 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B24B 1/00 20130101; F05D
2250/15 20130101; Y10T 29/49982 20150115; F01D 11/12 20130101; F05D
2250/73 20130101; F05D 2250/712 20130101; B24B 19/26 20130101; F05D
2240/11 20130101 |
International
Class: |
B24B 1/00 20060101
B24B001/00; B24B 19/26 20060101 B24B019/26; F01D 11/12 20060101
F01D011/12 |
Claims
1. A blade outer air seal for a gas turbine engine having an engine
rotation centerline, comprising: a substrate having a first end and
a second end, wherein a blade within the engine rotates past the
first end and then past the second end when the engine is running;
and a coating applied to the substrate; wherein the substrate and
the coating define a first combined thickness at the first end and
a second combined thickness at the second end; wherein the first
combined thickness is different from the second combined thickness,
wherein the coating has a first thickness at the first end and a
second thickness at the second end such that the coating varies in
thickness in a circumferential direction with respect to the engine
rotation centerline, and wherein the substrate has a substantially
uniform thickness.
2. The blade outer air seal of claim 1, wherein the coating
comprises a thermal barrier coating.
3. The blade outer air seal of claim 1, wherein a surface of the
substrate is eccentric with respect to the engine rotation
centerline.
4. The blade outer air seal of claim 1, wherein: a first surface of
the substrate is not eccentric with respect to the engine rotation
centerline; and a second surface of the coating is eccentric with
respect to the engine rotation centerline.
5. The blade outer air seal of claim 4, wherein the coating
comprises a thermal barrier coating.
6. The blade outer air seal of claim 1, wherein the coating has a
smooth transition from the first end to the second end.
7. A gas turbine engine having an engine rotation centerline,
comprising: a plurality of blade outer air seals mounted within the
engine, wherein each blade outer air seal comprises: a substrate;
and a coating applied to the substrate; wherein a surface of the
coating of each blade outer air seal is eccentric with respect to
the engine rotation centerline when the blade outer air seal is
mounted within the engine such that each coating has a first
thickness at a first end of a respective blade outer air seal and a
second thickness at a second end of the respective blade outer air
seal, and wherein the plurality of blade outer air seals form a
stair step configuration when mounted within the engine.
8. The gas turbine engine of claim 7, wherein the coating of each
blade outer air seal is a thermal barrier coating.
9. The gas turbine engine of claim 7, wherein the substrate of each
blade outer air seal is uniform in thickness.
10. The gas turbine engine of claim 7, wherein a surface of each
substrate of the plurality of blade outer air seals is eccentric
with respect to the engine rotation centerline.
11. The gas turbine engine of claim 7, wherein: a first surface of
the substrate is not eccentric with respect to the engine rotation
centerline; and a second surface of the coating is eccentric with
respect to the engine rotation centerline.
12. The gas turbine engine of claim 11, wherein the coating
comprises a thermal barrier coating.
13. The gas turbine engine of claim 7, wherein: wherein a blade
within the engine rotates past the first end and then past the
second end when the engine is running, the substrate and the
coating define a first combined thickness at the first end and a
second combined thickness at the second end, and the first combined
thickness is different from the second combined thickness.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation application of the
legally related U.S. application Ser. No. 14/176,669, filed Feb.
10, 2014, which claims the benefit of U.S. Provisional Application
Ser. No. 61/763,231, filed Feb. 11, 2013, the contents of which are
both incorporated by reference herein in their entireties.
TECHNICAL FIELD OF THE DISCLOSURE
[0002] The present disclosure generally related to turbine engines
and, more specifically, to a blade outer air seal of a turbine
engine.
BACKGROUND OF THE DISCLOSURE
[0003] Axial turbine engines generally include fan, compressor,
combustor and turbine sections positioned along an axial centerline
sometimes referred to as the engine's "axis of rotation" The fan,
compressor, and combustor sections add work to air (also referred
to as "core gas") flowing through the engine. The turbine extracts
work from the core gas to drive the fan and compressor sections.
The fan, compressor, and turbine sections each include a series of
stator and rotor assemblies. The stator assemblies, which do not
rotate (but may have variable pitch vanes), increase the efficiency
of the engine by guiding core gas flow into or out of the rotor
assemblies.
[0004] Each rotor assembly typically includes a plurality of blades
extending out from the circumference of a disk. Platforms extending
laterally outward from each blade collectively form an inner radial
flowpath boundary for core gas passing through the rotor assembly.
An outer case, including blade outer air seals (BOAS), provides the
outer radial flow path boundary. The blade outer air seal aligned
with a particular rotor assembly is suspended in close proximity to
the rotor blade tips to seal between the tips and the outer case.
The sealing provided by the blade outer air seal helps to maintain
core gas flow between rotor blades where the gas can be worked (or
have work extracted).
[0005] Disparate thermal growth between the rotor assembly and the
outer case can cause the rotor blade tips to "grow" radially and
interfere with the aligned blade outer air seal. In some
applications, the gap between the rotor blade tips and the blade
outer air seal is increased to avoid the interference. A person of
skill in the art will recognize, however, that increased gaps tend
to detrimentally effect the performance of the engine, thereby
limiting the value of this solution. In other applications, the
blade outer air seals comprise an abradable material and the blade
tips include an abrasive coating to encourage abrading of the blade
outer air seals. The blade tips abrade the blade outer air seal
until a customized clearance is left which minimizes leakage
between the rotor blade tips and the blade outer air seal.
[0006] Improvements are therefore needed in turbine engine rotor
assembly blade outer air seals that decrease the flow of core gas
around the rotor blade tips to increase turbine engine
efficiency.
SUMMARY OF THE DISCLOSURE
[0007] In one embodiment, a blade outer air seal for a gas turbine
engine having an engine rotation centerline is disclosed,
comprising: a substrate having a first end and a second end,
wherein a blade within the engine rotates past the first end and
then past the second end when the engine is running; a coating
applied to the substrate; wherein the substrate and the coating
define a first combined thickness at the first end and a second
combined thickness at the second end; wherein the first combined
thickness is selected from the group consisting of: greater than
and less than, the second combined thickness.
[0008] In another embodiment, a blade outer air seal for a gas
turbine engine having an engine rotation centerline is disclosed,
comprising: a substrate; and a coating applied to the substrate;
wherein a surface of the coating is eccentric with respect to the
engine rotation centerline when the blade outer air seal is mounted
within the engine.
[0009] In another embodiment, a method for creating a blade outer
air seal for a gas turbine engine having an engine rotation
centerline is disclosed, comprising the steps of: a) determining a
desired surface profile for the blade outer air seal; b) providing
a rotating grinding surface having a grinding rotation centerline;
c) determining where the engine rotation centerline would be if the
blade outer air seal were mounted in the engine; d) offsetting the
grinding rotation centerline from the engine rotation centerline;
and e) applying the rotating grinding surface to the blade outer
air seal while rotating the rotating grinding surface about the
grinding rotation centerline to create the desired surface
profile.
[0010] In another embodiment, a method for grinding a work piece
having nominal curvature defined by a work piece curvature
centerline is disclosed, comprising the steps of: a) determining a
desired surface profile for the work piece; b) providing a rotating
grinding surface having a grinding rotation centerline; c)
offsetting the grinding rotation centerline from the work piece
curvature centerline; and d) applying the rotating grinding surface
to the work piece while rotating the rotating grinding surface
about the grinding rotation centerline to create the desired
surface profile.
[0011] Other embodiments are also disclosed.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine.
[0013] FIG. 2 is a partial perspective view of a first stage high
pressure turbine blade and blade outer air seal showing an
inconsistent rub pattern.
[0014] FIGS. 3A-C are elevational views of a blade outer air seal
exhibiting a nonuniform coating thickness across its surface,
according to one disclosed embodiment.
[0015] FIG. 4 is a schematic elevational view illustrating an
eccentric grinding device and method according to one disclosed
embodiment.
[0016] FIG. 5 is a schematic elevational view of a series of blade
outer air seals, each having an eccentrically ground surface,
according to one disclosed embodiment.
DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENTS
[0017] For the purposes of promoting an understanding of the
principles of the invention, reference will now be made to certain
embodiments and specific language will be used to describe the
same. It will nevertheless be understood that no limitation of the
scope of the invention is thereby intended, and alterations and
modifications in the illustrated device, and further applications
of the principles of the invention as illustrated therein are
herein contemplated as would normally occur to one skilled in the
art to which the invention relates.
[0018] FIG. 1 illustrates a gas turbine engine 10 of a type
normally provided for use in a subsonic flight, generally
comprising in serial flow communication a fan 12 through which
ambient air is propelled, a compressor section 14 for pressurizing
the air, a combustor 16 in which the compressed air is mixed with
fuel and ignited for generating an annular stream of hot combustion
gases, and a turbine section 18 for extracting energy from the
combustion gases.
[0019] It has been observed in some turbine engines that the blades
of the first stage high pressure turbine create an inconsistent rub
on the blade outer air seal. Referring to FIG. 2, there is shown a
close-up view of a first stage high pressure turbine blade 100. As
is known in the art, gases flowing through the turbine engine
impact the blade 100, thereby causing rotation of the high pressure
turbine. The blade 100 moves away from the viewer in the view of
FIG. 2 when it is rotating.
[0020] The distal end 102 of the blade 100 is designed to rub
against the segmented blade outer air seal 104, thereby providing a
seal to prevent gases from flowing between the blade 100 and the
blade outer air seal 104. Energy that may be imparted to the
turbine is lost when such gases bypass the turbine blade, reducing
the efficiency of the engine. The area 106 of heavy rubbing on the
surface of the blade outer air seal 104 indicates consistent
contact with the distal end 102 of the blade 100 as it rotates by
the blade outer air seal 104, forming an effective seal
therebetween.
[0021] In some situations, portions of the blade outer air seal 104
may move farther away from the distal end 102 of the blade 100
during hot conditions of the engine. This may be caused by one or
more of a variety of causes, including heat, pressure, loads or
movement of adjoining hardware, etc. The area 108 of light and
inconsistent rubbing is indicative of this problem. Because the
distal end 102 of the blade 100 does not make consistent contact
with the blade outer air seal 104 in the region 108, energy that
would otherwise by transferred to the blade 100 is lost and the
efficiency of the turbine is decreased.
[0022] There is therefore a need for apparatuses and methods for
ensuring consistent contact between the distal end 102 of the blade
100 and the surface of the blade outer air seal 104. The presently
disclosed embodiments are directed toward solving this problem.
[0023] In the presently disclosed embodiments, methods are
disclosed for creating a non-uniform radial distance from the
centerline of a turbine engine to the inner surface of a static
piece of hardware, such as a first stage high pressure turbine
blade outer air seal. By varying this distance, it is possible to
promote substantially consistent rub between hardware rotating
around the engine centerline and static hardware positioned at a
nominal radial distance from the engine centerline Although the
concept is described herein with respect to rotating blades of a
first stage high pressure turbine and a segmented blade outer air
seal for such turbine, it will be appreciated from the present
disclosure that the disclosed concepts may be employed with any
system where it is desired to precisely control the contact (or
gap) between a piece of rotating hardware and a piece of static
hardware. For example, the presently disclosed concepts are also
applicable to any rotating hardware on a turbine engine where it is
desired to precisely control the contact (or gap) between the
rotating hardware and a piece of static hardware.
[0024] Referring now to FIG. 3A, one segment of a blade outer air
seal 200 according to one embodiment is illustrated in profile. The
blade outer air seal 200 consists of a main body 202 to which is
applied a thermal barrier coating 204, as is known in the art. It
is desired that the distal end 102 of the blade 100 maintain
consistent contact with the thermal barrier coating 204 as the
distal end 102 of the blade 100 moves across the surface of the
blade outer air seal 200.
[0025] In situations where it is observed that the distal end 102
of the blade 100 is not making consistent contact, such as in the
situation illustrated in FIG. 2, the seal may be repaired by
applying a second layer 206 to the thermal barrier coating 204. The
second layer 206 may comprise the same material as the thermal
barrier coating 204 or a different material, as desired. It can be
seen that at the end of the blade outer air seal 200 shown in
close-up in FIG. 3B, the second layer 206 is thicker than the
thickness of the second layer 206 shown in close-up in FIG. 3C at
the opposite end of the blade outer air seal 200. This causes a
total coating thickness of X in the portion shown in FIG. 3B and a
total coating thickness of Y in the portion shown in FIG. 3C, where
X>Y. The blade outer air seal 200 is thereby moved closer to the
distal end 102 of the blade 100 on the end of the blade outer air
seal 200 in the portion shown in FIG. 3B, and the thicker coating
206 will promote rub on the side that previously had reduced
contact, thereby closing the gap that was previously causing
inconsistent contact therebetween. It will be appreciated from the
present disclosure that the thicker coating thickness may be
located at any desired portion of the static hardware.
[0026] The differing thicknesses X and Y, as well as the smooth
transition therebetween (i.e., the desired surface profile), may be
created by grinding the second layer to an inconsistent thickness
across the width of the blade outer air seal 200. One embodiment
method for creating such a profile is illustrated schematically in
FIG. 4. A work piece, such as a blade outer air seal 200 to name
just one non-limiting example, may be ground by a rotating grinding
surface 300 that rotates about a grinding axis 302. The grinding
axis 302 may be moved in an arc 304 during the grinding process,
the arc having a grinding rotation centerline 306. The work piece
may have its own nominal curvature defined by a work piece
curvature centerline 308. For example, if the work piece is a blade
outer air seal 200 for use in a gas turbine engine having an engine
rotation centerline, the work piece curvature centerline 308
coincides with the engine rotation centerline (i.e., where the
engine rotation centerline would be if the blade outer air seal 200
were currently mounted within the engine). By offsetting the
grinding rotation centerline 306 from the engine rotation
centerline 308 by a distance 310, an eccentrically ground surface
will be created on the blade outer air seal 200.
[0027] Therefore, in one embodiment the method for creating the
eccentrically ground surface comprises the steps of: a) determining
a desired surface profile for the blade outer air seal 200; b)
providing a rotating grinding surface 300 having a grinding
rotation centerline 306; c) determining where the engine rotation
centerline 308 would be if the blade outer air seal 200 were
mounted in the engine; d) offsetting the grinding rotation
centerline 306 from the engine rotation centerline 308 by the
distance 310; and e) applying the rotating grinding surface 300 to
the blade outer air seal while rotating the rotating grinding
surface 300 about the grinding rotation centerline 306 to create
the desired surface profile.
[0028] The configuration and method discussed hereinabove with a
two layer (204 and 206) configuration is well-suited to repair
scenarios, as the existing structure is left intact and material is
added thereto and ground to the desired surface profile. In other
embodiments, the second layer 206 is omitted and the thermal
barrier coating 204 is subjected to the eccentric grinding process.
This is useful in applications where it is not required to keep a
uniform thickness to the thermal barrier coating. In other
embodiments, the ground substrate 202 (which is typically metal,
but may be formed from any desired material) is ground to the
desired shape, and then a uniform coating of the thermal barrier
coating 204 is applied thereto.
[0029] As shown in FIG. 5, a series of blade outer air seals 202,
each having an eccentrically ground surface, may be mounted within
a gas turbine engine. It can be seen that the thickness A on a
first end of the blade outer air seal 200 is greater than a
thickness B on a second end of the blade outer air seal 200. The
eccentric grind, either to the blade outer air seal substrate 202
or to the thermal barrier coating 204, on each of the blade outer
air seals 200 creates a stair step configuration when the blade
outer air seals 200 are mounted in the engine and are cold.
Choosing the proper eccentric profile will result in a circular
flowpath at the thermal barrier coating 204 surface in the running
engine when the blade outer air seals 200 are subjected to the
forces discussed above.
[0030] While the invention has been illustrated and described in
detail in the drawings and foregoing description, the same is to be
considered as illustrative and not restrictive in character, it
being understood that only certain embodiments have been shown and
described and that all changes and modifications that come within
the spirit of the invention are desired to be protected. For
example, those skilled in the art will recognize that in some
embodiments the work piece that is ground may be something other
than a blade outer air seal, as well as something other than a part
of a gas turbine engine. The disclosed concepts are applicable for
creating an eccentric profile on any type of workpiece.
* * * * *